Abstract: In this paper, the blunt leading-edge of a wing is taken as mesh points, and there the exact velocity potential equation with central difference scheme and the exact boundary condition are used, while in the other places, the approximate velocity potential equation, which assumes small perturbation in the transverse plane but allows large perturbation in the longitudinal direction, and the corresponding boundary condition are employed. Two numerical examples are following:（1） A rectangular wing having airfoil NACA0012, aspect ratio λ = 12, angle of attack α=2°, free stream Mach number M∞ = 0.63. The computed pressure distribution of the root section agrees with the exact numerical subsonic solution given by Sells （1968）.（ 2 ） The sweepback wing tested by NACA RM A51G31 having airfoil NA-CA64A010 which is perpendicular to 1/4 chord line with sweepback angle X1/4 = 45°, λ = 3 and taper ratio η = 2 ,α=2°, M∞=0.4, 0.8 and 0.9. The computed pressure distributions agree well with those obtained by tests.Under the assumption of local linearization, the stability of the difference equation in line relaxation with Seidel iteration is studied by the von Neumann method and the convergence of the solution of the differential equation equivalent to the above difference equation to the solution of the original differen- tial equation is discussed by the method of separation of variables.The following conclusions are obtained:（ 1 ） The stability condition for the line relaxation with Seidel iteration is 0<ω≤2 at locally subsonic points, where ω is the relaxation factor.（ 2 ） At locally supersonic points, the relaxation is always unstable. The convergence conditions are as follows. Let the steps Δx （chordwise） and Δz （spanwise） perpendicular to the relaxation line.（3） 0<ω<2, at locally subsonic points.（4） 0 <ω<1+ , at locally supersonic points.The numerical experiences agree with the conclusions （ 1 ）, （ 3 ） and （ 4 ）, but do not agree with the conclusion （ 2 ）.